Method for controlling the pitch attitude of a satellite by means of solar radiation pressure and satellite, in particular an electric propulsion satellite, suitable for implementation of the method

ABSTRACT

A satellite embodying at least one surface intended mainly for exposure to solar radiation and extending away from the satellite in a predetermined direction (Y), an on-board computer having connected thereto an attitude sensing system, an orbit control system for imparting thrust to the satellite along predetermined axes, and an attitude control system. The satellite further embodies a device for controlling the tilt of the surface in parallel with a plurality of planes containing the predetermined direction; and, therefore, particularly in parallel with the plane of a solar panel forming the surface. The tilt control device is controlled by the on-board computer. The tilting can generate a moment of pitch or relocate the center of gravity onto the axis of the orbit control system.

BACKGROUND OF THE INVENTION

1. Field of The Invention

The invention concerns controlling the attitude of a satellite about itsthree axes, usually a geostationary satellite stabilized about the threeaxes, with optional compensation of disturbing torques acting on thesatellite about the roll/yaw axes during orbit control maneuvers.

It is also concerned with the general configuration of satellites whoseattitude is stabilized about their three geometrical axes ingeostationary orbit, whether for civil or military, commercial orscientific purposes, or combinations thereof.

2. Description of the Prior Art

In the present context a satellite is any artificial object in the solarsystem in orbit around the Earth or any other planet or object of thesolar system, or, in solar orbit, possibly a transfer orbit between twoplanets.

Satellites in orbit are known to be subjected to disturbing torqueswhich make it necessary to control their attitude. The most importantcauses of disturbing torques are the lack of symmetry about the centerof gravity (in space the expression "center of mass" is moreappropriate) of the effects of solar radiation pressure (solar pressurefor short) due to the angle of incidence (not 90°) of the pitch axis ofthe satellite to the Sun, different reflectivity characteristics ofparts of the satellite and geometrical asymmetries of the satellite; theaction of the local (for example terrestrial) magnetic field; theaerodynamic effect of the environment (in low orbits); and the distancefrom the center of gravity of the satellite to the resultant thrustvector axis of the thrusters used to modify the Satellite orbit.

It is possible to distinguish between disturbances related to theenvironment: solar pressure, interaction of the satellite magneticdipole with the surrounding magnetic field, gravitational gradient,etc.; these disturbances are weak (order of magnitude=10⁻⁵ N.m) but acton the satellite at all times, and disturbances related to misalignmentof the thrust vector of the orbit control chemical thrusters relative tothe center of gravity of the satellite; these disturbances are stronger(order of magnitude=10 N.m.s per day for a geostationary satellite) butlimited in time.

It is essential to provide means for controlling the attitude of asatellite in its orbit. Various active means have already been proposedfor this purpose, using a plurality of reaction wheels or thrusters ofthe mass ejection type, but the principle of ejecting mass requires thatthe satellite carry a reserve of mass, which increases the weight of thesatellite at launch. Further, gas jet thrusters cause intensedisturbances which excite the flexible and nutation modes of thesatellite, degrading pointing accuracy. Additionally, low-thrust typethrusters such as ion thrusters or electrical arc ionization thrustersconsume considerable electrical power and require warm-up phases whichgenerally lead those skilled in the art to avoid using them for attitudecontrol, and reaction wheels are not sufficient in themselves becausethe wheels must be desaturated to bring their speed to a value near thenominal value from time to time and this requires the application ofexternal torque to the satellite.

To control the attitude of a satellite for an optimal mass budget use ismade of disturbing forces due to solar pressure, by appropriatelyorienting surfaces attached to the satellite, or the local, for exampleterrestrial, magnetic field, by creating magnetic dipoles on board thesatellite by means of pairs of currents.

Various prior art references have already put forward the use of solarradiation pressure for satellite attitude control and orbit control(stationkeeping) using mobile surfaces which can be oriented by means ofdedicated actuators or using orientation thrusters already on board.

French Patent 2,513,589 describes a method and a device for aligningwith a required direction the roll axis of a satellite which isspin-stabilized and fitted with a plurality of fixed solar panels;mobile surfaces are mounted at the ends of the panels.

French Patent 2,550,757 proposes to control the position of satellitesby acting on the solar panels by deforming them to impose a variablebackwards curvature on each of them.

French Patent 2,529,166 concerns a satellite stationkeeping method usingsolar sails and a space vehicle implementing this method. Solar sailsdisposed to the North and South are mounted on the satellite at the endof pylons parallel to the North-South axis. The pylons can rotate onthemselves and the sails can be inclined about axes transverse to thepylons.

German Patent 2,537,577 entitled "Satellite Attitude Control", teachesthe provision at the end of the solar panels of surfaces that can beoriented about the axis of the solar panels and transversely thereto.

U.S. Pat. No. 3,304,028 entitled "Attitude Control for Spacecraft", issimilar to French Patent 2,513,589, previously mentioned, as is U.S.Pat. No. 3,339,863.

French Patent 2,530,046 entitled "Geosynchronous Satellite AttitudeControl Method and Device", teaches the addition of fixed surfaces tothe sides of the solar panels.

French Patent 2,531,547 entitled "Geostationary Satellite AttitudeControl System", teaches variation of the relative orientations of thesolar panels about their axes as does U.S. Pat. No. 4,325,124 entitled"System for Controlling the Direction of the Momentum Vector of aGeosynchronous Satellite".

European Patent 0,295,978 proposes a device and a method for pointing aspace probe towards a heavenly body. North and South solar sails areadded to the satellite which have asymmetrical surface areas,orientations about a North-South axis or inclinations transverse to thisaxis.

French Patent 2,552,614 proposes a Satellite configuration with improvedsolar means comprising solar panels oriented transversely to theNorth-South axis and adapted to be oriented about axes transverse-to theNorth-South axis.

Finally, U.S. Pat. No. 4,262,867 provides for solar panels adapted to bepartially retracted accordion fashion to each side of which solar sailsare hinged about axes transverse to the longitudinal axis of the panels.

These prior art references concern attitude control devices which usesolar pressure as their means of actuation. However, all the solutionstaught therein have one or other of the following drawbacks. Either theyrequire extra surfaces to be added, with the disadvantages that theadditional surfaces increase the mass of the satellite; the addition ofmechanisms dedicated to deploying the surfaces in orbit increases themass and the risk of failure; and the additional overall dimensions dueto the surfaces represent a satellite volume penalty at launch; or theyprovide satellite attitude control about one or two axes only, requiringfurther means for control about the third axis.

Various documents have proposed displacing the center of gravity of thesatellite to reduce the disturbing torques related to misalignmentbetween the center of gravity and the thruster (or solar pressure)thrust vector. They include U.S. Pat. No. 4,684,084 entitled "SpacecraftStructure with Symmetrical Mass Center and Asymmetrical DeployableAppendages"; U.S. Pat. No. 4,345,728 entitled "Method for Controllingthe Attitude of a Spinning Spacecraft in the Presence of SolarPressure"; and U.S. Pat. No. 3,516,623 entitled "Stationkeeping System".

U.S. Pat. Nos. 3,516,623 and 4,345,728 propose reducing the disturbingtorques acting on a spin-stabilized satellite related to misalignment ofthe center of gravity and the thrust by moving the center of gravityusing mobile weights, these weights and their actuators having no otherfunction.

U.S. Pat. No. 4,684,084 describes a satellite configuration in which thedisturbing torques due to misalignment between the center of gravity andthe thrust vector of the orbit control thrusters are reduced. The centerof gravity is moved towards the thrust axis by appropriate positioningof the solar generator panels after they are deployed. This positioningis fixed and not variable in flight. This configuration is such that thecenter of gravity is substantially fixed despite the deployment ofhighly asymmetric appendages, but there is no possibility of modifyingthe position of the solar generator panels in flight. This has thedrawback of increasing the disturbing torques due to solar radiation andof providing nothing to compensate this.

As for the propulsion employed during the operational phase of currentthree-axis stabilized satellites, in particular in the United States,Japan and Europe, this is purely chemical (using hydrazine or a mixtureof propellants, for example) or chemical with electrical assistance(example: power augmented catalytic thruster (PACT) heated or electricalarc (Arcjet) hydrazine or ion or plasma thrusters for orbit correction.

However, in the final analysis attitude control is achieved by chemicalpropulsion with intermediate Biorage of angular momentum in one or moreinertia wheels about two or three axes.

Relevant publications include: "The Attitude Determination and ControlSubsystem of the Intelsat V Spacecraft"--Proceedings of the AOCSConference, Noordwijk, October 1977; "Precision Attitude Control with aSingle Body-Fixed Momentum Wheel"--AIAA Mechanics and Control FlightConference--Anaheim, Calif., August 1974; U.S. Pat. No. 4,949,922entitled "Satellite Control System"; and "Satellite Attitude and OrbitControl System: Developments to the 80-90's"--L'Aeronautique etl'Astronautique--no. 69, 1878-2--p 33-56.

Similarly, the use of electrical propulsion for orbit control and evenattitude control is under widespread consideration at the present time,as indicated by the following publications, in particular, "ElectricPropulsion Projects and Researches in Japan", AIAA 20th InternationalElectric Propulsion Conference, Garmisch, Partenkirchen (Germany),October 1988; "Design and Integration of an Electric Propulsion Systemon the Eurostar Spacecraft", same conference as above; "ReadinessAppraisal: Ion Propulsion for Communication Satellites", AIAA 12thInternational Communication Satellite Systems Conference, Crystal City,March 1988; and "Chemical and Electric Propulsion Tradeoffs forCommunication Satellites", Comsat Technical Review Volume 2 Number 1,Spring 1972, pp 123-145.

With reference to ion thruster propulsion as such, reference may be madeto French Patent 2,510,304 entitled "Field Emission Ion Source Suitablefor Electric Propulsion of Space Craft"; U.S. Pat. No. 3,279,176entitled "Ion Rocket Engine"; and U.S. Pat. No. 4,829,784 entitled"Method and System for Storing Inert Gas for Electrical Impulse SpaceDrives".

SUMMARY OF THE INVENTION

An object of the invention is to exploit the solar radiation pressureacting on preexisting surfaces provided primarily to be exposed to solarradiation (solar generator panels and/or solar sails) to control theattitude of a satellite in pitch reliably and simply, with the leastpossible additional mass, in a manner that is decoupled and compatiblewith any known type of roll/yaw control, also based on solar radiationpressure, for example.

Another object of the invention is to reduce the power requirement forattitude control, to minimize the attitude disturbance during orbitcontrol maneuvers with no significant mass or reliability penalty.

Another object of the invention is to enable redundancy to beincorporated into the solar generator panel drive motors which atpresent constitute a single point failure hazard.

Another object of the invention is to obtain maximum benefit forattitude control (about the three axes) and orbit control fromelectrical propulsion (the great advantage of which is a much betterspecific impulse than chemical propulsion), of a kinetic energy storagesystem advantageously with no gyroscopic stiffness based on reactionwheels (lighter in weight than inertia wheels which have a non-nullangular momentum at all times) and disturbing forces generated by thesolar radiation pressure, in order to be able to dispense with anychemical propulsion in the operational phase and to minimize the overallmass of the components of the satellite dedicated to attitude control(about three axes) and orbit control, at moderate cost (manufacture andlaunch) and with improved overall reliability (because of the eliminatedrisk of leakage associated with the use of chemical propulsion).

To this end the invention proposes a satellite having at least onesurface intended principally to be exposed to solar radiation andextending from the satellite in a given direction, an onboard computerand connected to the latter an attitude sensor system, orbit controlmeans adapted to apply to the satellite thrusts along given axes andattitude control means, characterized in that it further has means fortilting the surface parallel to a plurality of planes containing thegiven direction, the tilt means being controlled by the onboard computerwith or without action from the ground.

According to preferred features of the invention the surface intendedprincipally to be exposed to solar radiation is a plane solar generatorpanel extending along the given direction and connected to the satelliteby a drive motor adapted to rotate the panel about the given direction.

The tilt means include the drive motor and a second motor disposedbetween the drive motor and the solar generator panel. The second motoris a rotary motor having an axis inclined at a non-null angle α relativeto the given direction and the non-null angle is between 2° and 15°. Thetilt means include a pivot motor whose axis is transverse to the givendirection, which motor provides a range of movement of at most 15°relative to the given direction.

The tilt means include a second pivot motor whose axis is transverse tothe given direction and has a non-null inclination to the axis of thefirst pivot motor.

The tilt means include a linear motor extending in a direction inclinedto the given direction and mounted on one side of a deformablearticulated triangle coupling the surface of the satellite.

The orbit control means may be electric thrusters, or electric arcionization thrusters.

The attitude control means during orbit control maneuvers are thesurface and the tilt means or the attitude control means during orbitcontrol maneuvers is an orientable angular momentum system.

The invention also proposes a satellite adapted to be stabilized inattitude about roll, yaw and pitch axes in an at least approximatelycircular terrestrial orbit around the terrestrial North-South axis andincludes a body having North and South sides, an attitude sensor system,an onboard computer connected to the attitude sensor system, at leastone solar generator panel extending substantially parallel to the pitchaxis and coupled to the body by a device for rotating it about the pitchaxis under the control of the onboard computer so that the panel remainsat all times at least approximately perpendicular to the solarradiation, a kinetic energy storage system for at least three axescontrolled by the onboard computer and an attitude control and orbitcorrection propulsion system controlled by the onboard computer.

The attitude control and orbit correction propulsion system isexclusively electrical and has at least a first pair of two electricthrusters disposed substantially symmetrically relative to the plane ofthe pitch and yaw axes with non-null inclinations relative to the planeof the roll and yaw axes and to the plane of the pitch and yaw axes andinclinations of not more than approximately 20° to the plane of the rolland pitch axes.

The satellite further includes between the body and the solar generatorpanel a device for tilting the panel parallel to a plurality of planescontaining the pitch axis, the tilt device being controlled by theonboard computer.

According to optionally combinable preferred features of the inventionthe electric thrusters of the first pair are disposed near one of theNorth and South sides, near edges bordering the side parallel to the yawaxis, or are disposed substantially at the middle of the edges.

The propulsion system includes a second pair of electric thrustersdisposed substantially symmetrical to the plane of the pitch and yawaxes with non-null inclinations to the plane of the roll and yaw axesbut in the opposite direction to the thrusters of the first pair,non-null inclinations to the plane of the yaw and pitch axes andinclinations of not more than approximately 20° to the plane of the rolland pitch axes.

The first and second pairs of electric thrusters are substantiallysymmetrical to the plane of the roll and yaw axes.

The propulsion system has four electric thrusters only, and eachelectric thruster is inclined to the roll axis at an angle between 40°and 75° in absolute value and to the pitch axis at an angle between 15°and 65° in absolute value.

A second solar generator panel extends substantially parallel to thepitch axis on the opposite side to the first panel and is coupled to thebody by a second device for rotating it about the pitch axis under thecontrol of the onboard computer so that the second panel is at all timesat least approximately perpendicular to the solar radiation and by asecond device for tilting the second panel parallel to a plurality ofplanes containing the pitch axis, the panel tilt devices beingcontrolled by the onboard computer.

The kinetic energy storage system has no permanent gyroscopic stiffness,and at least three reaction wheels whose angular momentum can be reducedto zero.

This aspect of the invention is, therefore, not concerned either withthe use of electrical propulsion as such to carry out orbit correctionmaneuvers or with an attitude control system having no nominalgyroscopic stiffness as such, but a specific combination of the two,with the capability of displacing transversely to the pitch axis thecenter of thrust of the solar radiation pressure on the sails, so that:

1--it is never necessary to use chemical propulsion during theoperational phase (in practice the geostationary phase),

2--it is possible to provide only four electric thrusters (two may evensuffice if they are sufficiently reliable) to carry out orbitalcorrections in the North-South direction (approximately 50 m/s per year)and in the East-West direction (almost 5 m/s per year), even if onethruster fails, and

3--there is no requirement for a "gimbal", system for repointing theelectric thrusters.

With regard to the first point, the use of chemical propulsion duringthe geostationary phase is essential in all existing or currentlyplanned satellites for direct control of one to three axes of thesatellite using an attitude detector-processor-chemical actuatorfeedback loop;

to desaturate the reaction or inertia wheels which have stored angularmomentum about one to three axes of the Satellite (the overall feedbackloop is then: attitude detector-processor-angular momentum storage-speeddetector-chemical actuator); or

to meet the requirements of any axis by axis combination of the twofacilities (direct control or intermediate storage).

A satellite in accordance with this aspect of the invention useselectric propulsion to desaturate (reduce towards zero) the storedangular momentum about the yaw axis. Desaturation about the other axes(roll and pitch) is provided by conventional non-chemical means such assolar sails or magnetic loops. Desaturation of the stored angularmomentum about the yaw axis is rendered possible by the necessity tocarry out at least one orbit correction maneuver each day. It issufficient to carry out this maneuver in such a way that one of theelectrical thrusters operates for longer than the other.

Electrical propulsion with a thrust of 10 to 40 mN (millinewtons) perthruster requires daily maneuvers of almost one hour's duration,compared with twice-weekly maneuvers of one hundred seconds' duration,typically for thrusts in the order of 10 Newtons, with chemicalpropulsion.

For equal lever arms, the disturbing moments accumulated during atypical North-South maneuver are therefore around one tenth forelectrical propulsion as compared with chemical propulsion and can beaccumulated in reaction wheels, for example, without requiring largerinertia type wheels. On the other hand, in the case of chemicalpropulsion the disturbing torques induced by orbit correction maneuversare so high that kinetic or reaction wheels are unable to compensate forthem.

Only electric propulsion allows the use of small reaction wheels forabsorbing disturbing moments and, therefore, preserving the attitude ofthe satellite. By definition, small reaction wheels do not impart anynominal gyroscopic stiffness to the satellite. The mean value over thelife of the satellite of the accumulated angular momentum, axis by axis,is zero or very near zero, with a wide range of variation either side ofzero.

With regard to the second point above, the use of only four thrusters toprovide for inclination corrections (North-South) and drift corrections(East-West), even with one thruster failed, is also based on thejudicious use of electrical propulsion because its specific impulse isvery much higher than that of chemical propulsion (1,500 to 3000 sagainst typically 300 s), the propellant budget for drift correctionbecomes negligible (typical 40 to 50 kg against 500 for chemicalpropulsion) and it is possible to use a single thruster instead of two.If the thrusters are appropriately oriented, the resulting disturbingtorques can be sorted temporarily in small reaction wheels which aredesaturated later using low external torques (solar sails, magneticcoils).

The possibility of tilting the panels relative to the pitch axis has thefollowing advantages.

Failure of any one of the four thrusters will not endanger the mission.on current chemical systems the loss of any one of at least twelvethrusters means that one entire branch of the system (six thrusters)must be shut down and only the other, redundant branch used. A secondfailure terminates the mission if chemical propulsion is used but merelyreduces its duration if electrical propulsion is used.

With reference to the third point above, recently developed concepts ofusing electrical propulsion on board geostationary satellites (the ESAARTEMIS satellite, for example) require mounting the electricalthrusters on gimbals so that they can be oriented optimally to reducethe disturbing torques during maneuvers.

These satellites use chemical propulsion to eliminate the accumulatedangular momentum. As the thrust developed by chemical thrusters isrelatively high, the pulses must be very short to avoid disturbing theattitude of the satellite. This reduces the overall efficiency of theoperation because the specific impulse of the chemical thrusters usedwith very short pulses (in the order of ten milliseconds) is very low.To avoid the need for prohibitive quantities of propellants, an attemptis made to eliminate the primary cause, in other words the lever arms ofthe disturbing torques, by reorienting the electrical thrusters inflight.

Our concept of a satellite with no nominal gyroscopic stiffness, totallyunsuited to chemical propulsion during orbit control maneuvers, issuited to the slow accumulation of angular momentum during maneuvers andto a slow return towards a virtually null global angular momentum by theaction of low external torques (solar sails, magnetic torques).

According to other optionally combinable preferred features of theinvention:

at least some of the electric thrusters are ion thrusters;

at least some of the electric thrusters are plasma thrusters;

a system for attitude control during the phase of injection into theoperational orbit is connected to the same single-propellant storagetank as the electric thrusters;

the satellite constitutes a dual propulsion and attitude controlpropellant system for the operational orbit injection phase;

the rotation device and the tilt device of each panel are controlled bythe onboard computer to generate attitude correction about the threeaxes using the solar radiation pressure on each panel; and

the onboard computer is adapted to control yaw attitude correction bythe electric thrusters.

The invention also proposes a method of controlling the attitude of asatellite having a surface intended principally to be exposed to solarradiation and extending from the satellite in a given direction, anattitude measuring device and a computer adapted to determine the valueof an attitude correction torque to be applied in pitch, characterizedin that the surface is tilted transversely to the solar radiation so asto generate a pitch torque substantially equal to the required attitudecorrection torque.

According to other preferred features of the invention, prior toapplying an orbit control thrust in a given direction the amplitude andthe direction of the offset of the satellite center of gravity from thisgiven thrust direction resulting from attitude disturbances during aprevious orbit maneuver are estimated and the surface is tilted at leastin part parallel to a plane containing the given direction in which thesurface extends and the direction of the difference so as to move thecenter of gravity towards the given thrust direction. During each orbitcontrol maneuver the attitude of the satellite is stabilized by actionof an orientable angular momentum system after which, before the nextorbit control maneuver, the surface is tilted parallel to the solarradiation to stabilize the satellite in pitch and to move the orientableangular momentum into a given orientation relative to the satellite.

Thus, in a method in accordance with the invention, the position of thecenter of the solar panel surface is moved laterally, by movement intranslation transversely to the pitch axis. As a result, the solarpressure generates a torque in pitch which enables attitude control,wheel speed desaturation and/or pitch disturbing torque compensation.

The solar generator may be displaced by any appropriate known actuatorswhich are not in themselves part of the present invention.

The benefit of the invention lies in the following advantages.

The pitch control obtained does not modify, or modifies only minimally,the roll/yaw behavior of the satellite, so that it can be combined withany known type of roll/yaw control; solar control using panel drivemotors is particularly indicated (see in particular French patentapplications 89-15732 and 89-17479 respectively filed 29 Nov. and 29Dec. 1989).

The lateral displacement of the solar panels displaces the center II ofgravity of the satellite enabling it to be aligned with the thrust ofthe North, South, East and West orbit control thrusters; this reducesvery considerably roll/yaw disturbance during North-South orbit controlmaneuvers and pitch disturbance during East-West maneuvers.

The additional mass of the actuators is compensated by the saving inpropellant achieved by dispensing with ejection of material for attitudecontrol (at least in pitch) and by minimizing roll/yaw disturbanceduring orbit control maneuvers.

The additional actuators can be used to back-up the conventional drivemotors for rotating the solar generator panels if the configurationadopted is one in which a motor with a slightly inclined axis is stackedon the drive motor; this redundancy is a substantial benefit overconventional solutions in which the drive motors constitute a singlepoint failure hazard.

The use of this concept makes it possible to design a satellite in whichthe use of gas jet thrusters would be abandoned during orbit, forexample in favor of electrical thrusters (ion, electric arc ionizationor plasma type) dedicated to orbit control only. A concept of this kindis advantageous because of the pointing accuracy permitted by theabsence of disturbances due to the thruster gas jets, and because of thesaving in propellant mass made possible by the better specific impulseof electrical thrusters as compared with gas jet thrusters.

Objects, features and advantages of the invention will emerge from thefollowing description given by way of non-limiting example withreference to the appended drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagrammatic view of a spacecraft stabilized about its threeaxes in orbit about the Earth;

FIG. 2 is a block diagram of attitude control logic for a satellite inaccordance with the invention incorporating by way of example the use ofroll/yaw control as proposed in French patent 2,655,167 and the use ofangular momentum, each of the three components of which can be varied;

FIG. 2A is a block diagram of component parts of the control logic fromFIG. 2;

FIG. 3 is a diagrammatic perspective view of the satellite in aconfiguration in which the solar pressure generates a pitch torque aboutthe Y axis of the satellite;

FIGS. 4, 5 and 6 show three known wheel arrangements adapted (to enableuse of the logic shown in FIG. 2) to generate angular momentum, allthree components of which can be varied;

FIG. 7 is a detail view showing how the movement of a solar panel of asatellite in accordance with the invention is controlled, using a rotarymotor whose shaft is slightly inclined to the panel drive motor;

FIG. 8 shows a variation of this arrangement using a pivoted 1. drivemotor with its axis transverse to the axis of the drive motor;

FIG. 9 shows another variation of this arrangement using an articulateddeformable triangle to couple the solar panel to the drive motor;

FIG. 10 is a diagrammatic view of the ideal configuration of an idealsatellite during an orbit control maneuver;

FIG. 11 shows a real configuration of a real Satellite;

FIG. 12 shows a real configuration of the real satellite improved inaccordance with the invention;

FIG. 13 is a perspective view to a larger scale of the body of thesatellite from FIGS. 1 and 3;

FIG. 14 is a perspective view of the satellite body as seen in FIG. 13rotated 180°;

FIG. 15 is a diagram of the thruster propellant supply circuit; and

FIG. 16 is a block diagram of the control system for the satellite fromFIGS. 13 and 14.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 shows a satellite 1 in a circular Earth orbit 2, for example ageostationary orbit inclined at not more than 5°, for example at 0.5°.

The satellite has a central body 3 with which are associated three axesdefining a satellite-related direct frame of reference. An axis Xtangential to the orbit 2 and in the same direction as the orbitalspeed, conventionally called the roll axis; an axis Y perpendicular tothe plane of the orbit 2 and oriented in the Earth's North-Southdirection, conventionally called the pitch axis; and an axis Zperpendicular to the X and Y axes and directed towards the Earth,conventionally called the yaw axis.

On some satellites this frame of reference may be oriented differentlyto the orbit and/or the Earth depending on mission-related imperatives.

The satellite has an attitude control device, as will be explainedbelow, by which it is stabilized about its three axes.

The satellite includes an attitude measuring device (conventional initself) connected to a processor (also conventional) which calculatesthe corrections to be applied by the attitude control device. The deviceconventionally includes (see FIG. 3) an Earth sensor 7, for example ofthe infrared type, usually adapted to measure the attitude in roll andin pitch. However, it may instead include Sun sensors or star sensors(now shown) to measure the attitude, in particular the attitude in yaw,if required. The processor is in practice part of an onboard computer 8shown in dashed line in FIG. 3.

Also conventionally, the satellite includes a solar generator with twopanels 4 and 5 extending, respectively, towards the North and towardsthe South along longitudinal axes coincident with the Y axis. The panelscan be oriented relative to the central body 3 about rotation axes whichare at least approximately coincident with the Y axis by two separatelycontrolled drive motors of any appropriate known type and one of which,mounted on the North side of the satellite, is denoted by the referencenumber 6 in the combination 6+6' of FIG. 3; these motors are normallyintended to hold the panels substantially facing the Sun,perpendicularly to its rays.

On some satellites the solar generator has only a single panel carriedon the North and South side. In this case the satellite may also carry asolar sail on the opposite side (either orientable or symmetrical aboutan axis parallel to Y) the function of which is to rebalance thesatellite as a whole with respect to the position of the center ofgravity and the position of the mean thrust due to solar pressure. Thistype of configuration does not rule out the use of the presentinvention. There can even be more than two panels or sails parallel tothe pitch axis.

Herein the term "solar generator" refers to the combination of one ortwo (or even more) panels and the term "solar generator panel" denotesassemblies which can be oriented by drive motors, in other wordsassemblies constituting the solar generator itself, that is to say theset of cells converting the light energy into electrical energy by thephotovoltaic effect, for example; the structure supporting thesecomponents; the mechanisms coupled to the structure to enable it to bestowed before the satellite reaches its orbital position, to deploy itand to hold it in the deployed position; and all the additionalcomponents which, in the orbital configuration, are fixed to thestructure and which have various functions, for example thermalprotection flaps which limit heat losses from the satellite duringphases in which the solar generator is not deployed or surfaces forincreasing the light impinging on the photovoltaic devices (shadowuniformization screens, for example).

In some cases deployable heatsinks fitted to the satellite can be usedas surfaces exposed to the solar pressure.

In practice the satellite also includes various appendages (antennas,etc.) which are fixed or virtually fixed and whose exposure no the solarpressure causes disturbing torques which are combined with thoseresulting from any asymmetry of the solar generator. Telecommunicationsatellites, for example, usually include one or more transmissionantennas and the radiation pressure of the transmission beam generates adisturbing torque about the Y axis which is added to those previouslymentioned.

In a manner that is also conventional, the satellite has orbit controlthrusters 9 of the chemical type, for example, the function of which isto return the satellite to its nominal position in space at regularintervals. Orbit control thrusters are needed because of the tendency ofany orbiting object to be moved out of its initial orbit by variousdisturbing forces. For example, in geostationary satellites theattraction of the Moon and Sun causes unwanted inclination of the orbitand the anisotropic shape of the Earth causes a drift towards the Eastor the West of the apparent position of the satellite relative to theEarth. In all cases this system has become necessary to modify the orbitof the satellite when this is a requirement of the mission, even if onlyat the end of its useful life, or injection into a cemetary orbit.

According to another aspect of the invention to be explained withreference to FIGS. 13 and 14 the system utilizes only electricthrusters; this will be explained later.

The satellite 1 is provided with a pitch attitude control systemcompatible with any other known means of satellite control in roll andyaw , in particular solar control, for example of the type described inone of the following references, namely, French Patents 2,655,167,2,656,586, 2,531,547, or French Patent 2,530,046, or U.S. Pat. No.4,325,124. As explained below, this pitch control is decoupled fromroll/yaw attitude control.

Also, the present invention is compatible with the philosophy underlyingthe teachings of the references mentioned above, hereby incorporated byway of reference, which is to add only minimum items to the satellite,or even none at all.

Coupled to a device of the same kind as those mentioned above, theinvention makes it possible to use solar radiation pressure to controlthe attitude of the satellite about its three axes.

Referring to FIG. 3, the pitch control torque is obtained by controlledtilting transversely to the solar radiation of at least one of the twosolar generator panels (and/or the solar sail if the satellite has arebalancing sail). This displaces the center of solar thrust relative tothe pitch axis Y and so generates a pitch torque.

During orbit control maneuvers this tilting can also be used to displacethe center of gravity of the panels and therefore the center of gravityof the satellite as a whole. The center of gravity of the satellite cantherefore be moved onto the axis of the thrust vector of the North-Southor even East-West control thrusters 9 which minimizes (or even cancels)roll and yaw disturbances due to orbit control (see below with referenceto FIGS. 10, 11 and 12). In practice, these disturbing torques are notalways totally eliminated but they are at least strongly attenuated sothat roll/yaw attitude control during these maneuvers can be entrustedto actuators (such as reaction wheels) which are less powerful andtherefore less costly in terms of mass than inertia wheels and whichexcite to a lesser degree the flexible and nutation modes of thesatellite; pointing accuracy will be improved accordingly.

The positioning of the center of gravity may be commanded on the basisof the identification (either automatic or by analysis on the ground) ofdisturbances during previous maneuvers. During these maneuvers the pitchtorque generated by the solar pressure is obviously negligible incomparison with the disturbance torques generated by the orbitcorrection maneuvers and the action of the actuators intended tocompensate for them; pitch control by tilting therefore ceases to beoperational, all the more so in that displacement of the center ofgravity towards the axis of the thrust vector can, in some cases,introduce an additional disturbing torque rather than correct theattitude in pitch. Another means of pitch control must therefore beused, for example one varying the speed of a wheel which is desaturatedsubsequently pending the next orbit control maneuver.

This is shown in FIG. 2 which schematically represents the two modeswhich alternate.

In orbit control mode (frame A) the center of gravity is displaced fromits estimated position after the previous maneuver towards the thrustvector axis of the orbit control thrusters. Then, on the basis ofinstantaneous attitude measurements, a variable angular momentum system9' (see FIG. 2A) is operated while the orbit control thrust is applied.

FIGS. 4 through 6 show various known variable angular momentum systemsthat can be used for the system 9. They include either a pitch kineticwheel 10 and two reaction wheels 11 and 12 with two axes in the roll/yawplane, or a pitch kinetic wheel 13 mounted on a double pivot 14, orthree reaction wheels 15, 16 and 17 on the pitch axis and two axes inthe roll/yaw plane. The axes chosen in the roll/yaw plane may of coursebe the roll and yaw axes. This latter embodiment has no gyroscopicstiffness.

In attitude control mode (frame B), tilting of the panels is commandedon the basis of attitude measurements, estimated disturbing torques andthe measured angular momentum still to be compensated after the previousorbit control maneuver in order to maintain the attitude of thesatellite and to desaturate the wheels.

The implementation of the control logic from FIG. 2 for determining theamplitude of the tilting to be applied to the panels will be obvious tothose skilled in the art.

FIG. 2A is a diagrammatic representation of the components of thecontrol logic: the onboard computer 8 operating on the drive motors 6,the tilt motor(s) 6', the orientable variable angular momentum system 9'and the orbit control thrusters 9, using various signals produced inknown ways by attitude sensors (including the sensor 7), from theorientable angular momentum System 9' or even received from othercomponents of the satellite or transmitted from the Earth.

The foregoing description considers attitude in pitch as well as in rolland yaw; roll and yaw control are advantageously effected as disclosedin French reference No. 2,656,167 already mentioned and herebyincorporated by way of reference.

One way among others to tilt the panel or sail parallel to its plane isto use a motor 6 is shown in any of FIGS. 7 through 9.

An orientation motor 20 whose axis is perpendicular to the plane of thepanels may be fitted between the conventional panel drive motor 6 andthe panel mounting arm 21 (see FIG. 8); the range of angular movement isfor example ±5° either side of the axis of the motor 6; the figure showspart of the cable 22 conveying power from the solar generator; or

a linear motor 30 (of the recirculating ball lead screw and rack type,for example) mounted on one or more of the solar generator mounting arms31 (Bee FIG. 9); this arm is then deformable; or

the preferred solution, adding a second motor 40 which may be identicalto the panel drive motor; the axis of this motor is slightly inclined atan angle a of around 5° for example to that of the drive motor (see FIG.7); in this case the panel (part of a mounting arm 41 which is shown inthe figure) is tilted by the two motors conjointly, the longitudinalaxis of the panel sweeping out a cone around the rotation axis of thepanel drive motor 6 while continuing to face in substantially the samedirection.

This latter solution has many advantages, including:

with only one additional motor it enables displacement of the center ofgravity in any direction in a plane parallel to the XZ plane, whichprovides not only for aligning the center of gravity on the North-Southaxis for the North-South orbit control maneuver but also for moving thecenter of gravity towards the East-West axis for East-West orbit controlorbit;

it is accommodated easily in the body of the satellite, facilitating thearrangement of the solar generator which does not need to be raised toprovide room for the additional mechanisms; and

should any of the conventional solar generator drive motors fail,redundancy is provided by the addition of the motor with the slightlyinclined axis; should this occur, pitch control and center of gravitypositioning are reduced in effectiveness by 50%; however, the satelliteis spared the serious consequences of loss of one drive motor onconventional satellites.

The axes X_(s), Y_(s) and Z_(s) are the axes of the frame of referencerelated to the satellite.

FIG. 10 shows the ideal case in which the center of gravity 0 is exactlywhere required and the thrust vector axis of North-South thrusters 50and 51 passes through this required position. In this ideal case aNorth-South orbit control maneuver does not directly induce anydisturbing torque.

In reality the center of gravity 0' is offset relative to the requiredposition and the thrust vector axis of the thrusters 50 and 51 does notpass exactly through this required position (FIG. 11). To minimize thedisturbing torque the invention teaches displacement of the real centerof gravity towards the thrust axis (position 01, in FIG. 12).

It is, therefore, possible on a satellite in accordance with the presentinvention to dispense with the use of conventional gas jet thrusters forattitude control on station apart from orbit control maneuvers, andduring orbit control maneuvers, the reduced disturbances enablingattitude control by less powerful means such as kinetic or reactionwheels, for example. The angular momentum stored in these wheels is thenvery easily desaturated by solar pressure attitude control whichoperates apart from orbit control maneuvers.

It will be appreciated that:

a) the tilting of one or more panels of the solar generator (or of thesail opposite the solar generator if the latter is asymmetrical) can beused either to create a pitch control torque by the effect of the solarradiation pressure on the solar generator panels (or the solar sail) orduring orbit control maneuvers to position the center of gravity of thesatellite in such a way as to minimize the disturbances due to thediscrepancy between this center of gravity and the real thrust vector ofthe orbit control thrusters, the two applications being usableseparately or together on the same satellite;

b) the tilting obtained by any known type of actuator enables use of thesolar pressure on the panels of the solar generator for roll and yawcontrol by any known method (which is not part of the presentinvention);

c) tilting (up to ±15°) can be achieved in two directions of theroll/yaw plane if a rotation actuator is selected whose axis is slightlyinclined to the axis of the solar generator panel drive motor. Thetilting is achieved by the differential rotation of the two motors: thedrive motor and the additional motor;

d) the angle between the axes of the two motors is between 2° and 15°;

e) the second motor can be used as back-up for the panel drive motorshould it fail;

f) the tilting (up to ±15°) can be obtained in one direction by a linearactuator mounted on one or more mounting arms of the solar generatorpanel;

g) tilting can also be achieved in one direction by a rotary actuatorwhose axis is at least approximately perpendicular to the plane of thesolar generator panels;

h) in an embodiment that is not shown tilting may be obtained in twodirections by two rotary actuators with different axes at leastapproximately perpendicular to the axis of the solar generator paneldrive motor;

i) attitude control about at least one of the roll, pitch and yaw I axesis provided totally or in part by a system utilizing solar pressure onsurfaces of the satellite;

j) in an embodiment that is not shown attitude control about at leastone of the roll, pitch and yaw axes is provided totally or partially bya system utilizing a magnetic dipole onboard the satellite;

k) attitude control about at least one of the roll, pitch and yaw axesis provided totally or partially by a system utilizing the variation ofat least one of the components of the angular momentum onboard thesatellite;

l) attitude control during orbit control maneuvers is obtained byvarying the three components of the onboard angular momentum and byminimizing disturbances by means of the invention;

m) attitude control outside orbit control maneuvers is obtained usingsolar pressure (in an embodiment that is not shown with the assistanceof magnetic dipoles);

n) in an embodiment that is not shown attitude control outside orbitcontrol maneuvers is provided by magnetic dipoles, possibly with theassistance of solar pressure;

o) attitude control outside orbit control maneuvers is used to align thecomponents of the onboard angular momentum with a set point value;

p) the satellite is a geostationary satellite;

q) all or some of the orbit control thrusters are of the low-thrust type(<1 Newton);

r) all or some of the orbit control thrusters are of the ionic type

s) all or some of the orbit control thrusters are of the electric arcionization type; and

t) all or some of the orbit control actuators are orientable surfacesexposed to the solar pressure.

The remainder of the description, which refers to FIG. 13 through 16,concerns an application wherein the orbit and attitude controlpropulsion system used in the operational phase does not use chemicalpropulsion. It is, therefore, free of the drawbacks inherent to chemicalpropulsion, such as risk of leaks, sudden variations in attitude and/orpossible generation of vibrations. Instead, the excellent specificimpulse of electrical thrusters (5 to 10 times greater than conventionalchemical thrusters) is used.

In a minimal configuration an electrical propulsion system 60 includesonly two thrusters 61 and 62 (or 63 and 64) disposed substantiallysymmetrically relative to the plane of the Y and Z axes, with aninclination al of approximately 40° to the pitch axis. This angle αl isusually between 15° and 65° in absolute value (meaning that its cosineis between 0.43 and 0.97 in absolute value) and the cosine of theinclination α2 of these thrusters to the roll axis is between 0.25 and0.75 in absolute value (angle between 40° and 75° in absolute value).

These thrusters are preferably perpendicular to the yaw axis.Constraints of available space may require that these thrusters areinclined at an angle α3 to the yaw axis between 70° and 110° (cosinebetween -0.35 and 0.35).

The thrusters are advantageously disposed near the edges of thesatellite body around a common side of the satellite (in this instancethe North side for the thrusters 61 and 62) parallel to the Z axis; thethrusters are preferably at the middle of these edges.

In the absence of any failures, a pair of thrusters is sufficient forEast-West orbit corrections (thrusts are generated parallel to the Yaxis at the same time, it is true) and orbit corrections parallel to theY axis. The requirement for a thrust towards the South is achieved by athrust towards the North one half-orbit later.

The propulsion system 60 advantageously includes a second pair toelectric thrusters 63 and 64 disposed near the opposite side (the Southside), each substantially symmetrical to the other about the YZ plane;the two pairs of thrusters are preferably symmetrical to each otherabout the XZ plane.

A propulsion system of this kind with only four electric thrusters 61through 64 provides for all necessary orbit correction operations (usingthese thrusters in pairs as much as possible), even should one thrusterfail.

Without departing from the scope of the invention it is possible to usea greater number of electric thrusters to increase reliability.

In principle the orientation of the thrusters 61 through 64 relative tothe body is fixed. In a more sophisticated version, however, they may berendered orientable, although this increases the mass and reducesreliability.

The satellite also includes a kinetic energy storage system with nogyroscopic stiffness formed by the reaction wheels 15 through 17 fromFIG. 6 which are used to temporarily store an overall angular momentumabout any axis. Speed sensors 15A, 16A and 17A are, in practice,associated with the reaction wheels and, in particular, are adapted toprovide a pitch angular speed signal.

The satellite in its transfer orbit (pending transfer to its operationalorbit) is in a configuration and an orientation different than those itassumes thereafter, which requires a dedicated attitude control system(not shown in FIG. 13 and 14) for this injection phase. Referring toFIG. 15, this system 70 advantageously uses thrusters 71 using the samesingle-propellant as the electric thrusters of the system 60. Thevarious thrusters are supplied from a storage tank 72 via pressureregulator and flow rate regulator systems 73 and 74 of any appropriateknown type.

In an embodiment of the satellite that is not shown, the satelliteincludes a dual propellant system for propulsion and attitude controlduring the phase of injection into the operational orbit, geostationaryor otherwise.

In the example shown (see FIG. 14) the satellite has on its -Z side apropulsion system 80 for injecting it into geostationary orbit; this maybe a solid fuel system.

The electrical propulsion system 60 can contribute to the injection ofthe satellite into its operational orbit, geostationary or otherwise.

For the injection phase, the satellite depicted in FIG. 14 may furtherinclude an additional electric propulsion system composed of twoelectric thrusters 90 parallel to the -Z axis.

The temporary reduction to zero of the angular momentum accumulated bythe reaction wheels in the operational orbit, geostationary orotherwise, is preferably carried out with respect to the three axes X, Yand Z using the solar radiation pressure on the panels in combinationwith the action of the thrusters.

Alternatively, attitude correction in roll and in yaw may be obtainedusing magnetic loops interacting with the terrestrial magnetic field.

Of course, the reduction to zero of the instantaneous angular momentumcomponent is easily achieved by means of the electric thrusters 61through 64, all that is required is a different period of operation oftwo thrusters required to operate simultaneously.

Although in the foregoing description it has been regarded asparticularly beneficial to have no gyroscopic stiffness, it should beunderstood that the invention is generally applicable to the case of anangular momentum having a continuously non-null component, for example acomponent along the Y axis (and therefore with an inertia wheel having acontinuously non-null angular momentum about the Y axis, as in FIGS. 4and 5, for example).

The number of wheels may advantageously be greater than three to provideredundancy.

For example, the satellite shown is a telecommunication satelliteweighing 3,000 kg at launch with large (80 m²) solar generators designedto generate 10 kill after 15 years.

Four ion thrusters (such as those supplied by MESSERSCHMITT-BOLKOW-BLOHM(MBB)) with a thrust of 100 millinewtons are disposed in the XY planewith a 60° slant relative to the Y axis. This configuration has theadvantage of good efficiency for the North and South thrusts and alsomakes it possible to limit the disturbing torque about the Z axis to anacceptable value for the reaction wheels in the event of degradedoperation with one thruster failed. The three reaction wheels have acapacity of ±15 Nms. They use friction-free magnetic bearings and onlythe electrical part is redundant, inside each wheel.

The solar generators are pointed towards the Sun at all times by thedevice 6+6' which rotates the generator relative to the satellite onceeach day and also enables the longitudinal axis of the solar generatorto be inclined a few degrees (70 for example) to the Y axis of thesatellite. Controlled by the onboard computer, these two movements areused to desaturate the reaction wheels, in other words to slow them downby generating the necessary solar torque. Solar control applies to thethree axes of the satellite.

Orbit maneuvers take place twice each day, lasting about one hour, therequired electrical power of approximately 1.5 kill being provided by abattery which is charged between two maneuvers.

This fine control concept (no angular momentum, no chemical thrusters)is such that any failure can only cause very slow drift of the satellitewhich is easily observable and quickly compensated by switching to theback-up unit for the failed unit. The periods of depointing are,therefore, minimized and the mission function is guaranteed at alltimes, which is a fundamental advantage.

A set of chemical thrusters is used nominally for the transfer orbitalone. Four chemical thrusters are sufficient but eight thrusters (notshown) are required to guard against failure. They are grouped togetheron the side away from the Earth around the apogee thruster. After thefirst few weeks of satellite operation these thrusters are isolated bysolenoid valves which eliminates all risk of leaks and is ofconsiderable advantage because no emergency action is required bycontrol stations, always difficult in the event of propellant leaks;there is no risk of thermal changes due to the consequences ofevaporation if leaks occur; and there is no risk of reduced missionduration following propellant leaks.

In exceptional cases of multiple failures there is provision for openingthe solenoid valves to point the satellite towards the Sun in so-calledsurvival mode pending expert advice on resuming solar control, with thechemical thrusters shut off again.

A mass balance associated with the four thrusters and their fuel, ascompared with that of a conventional system with 12 chemical thrusters,shows a saving of around 800 kg. For a satellite with a launch weight offour tons and a mission life of 15 years, the additional dry mass is 70kg for the electric thrusters but the fuel saving [(chemicalpropulsion)-(Xenon propulsion)] is 900 kg.

FIG. 16 is a block diagram of the control system. It is very similar toFIGS. 2 and 2A.

It will be realized that the invention proposes a novel combination ofcomponents known in themselves and already proven in orbit over a periodof many years, such as magnetic bearing wheels (SPOT satellites), andsolar generator rotation devices (all geosynchronous satellites).

It goes without saying that the foregoing description has been given byway of non-limiting example only and that numerous variants may beproposed by one skilled in the art without departing from the scope ofthe invention. The invention applies to any satellite having at leastone surface intended principally to be exposed to solar radiation andextending from the satellite in a given direction. The order of thedrive and tilt motors may be reversed, the drive motor being disposedbetween the tilt motor and the solar generator panel. The range ofmovement of the tilt motors may be increased without altering theprinciple of the invention.

The invention also applies to any satellite in respect of which thenecessary calculations are carried out in whole or in part on theground.

We claim:
 1. An attitude control system for a satellite, said attitudecontrol system comprising:a satellite having attitude stabilizing meansfor controlling said attitude about a pitch, yaw and roll axis whilesaid satellite is in a geostationary orbit; an onboard computer mountedto said satellite, said attitude stabilizing means being incommunication with said onboard computer; at least one surface memberextending from said satellite in a first direction away from saidsatellite, said at least one surface member adapted to be exposed tosolar radiation; orbit control means mounted to said satellite forapplying to said satellite thrusts along said pitch, yaw and roll axis;said orbit control means further comprising thruster means symmetricallyspaced about said pitch and yaw axis; said thruster means being inclinedat a predetermined angle at least with respect to said pitch axis suchthat a resultant force vector of said thrusts generated by said thrustermeans is in a first resultant force direction; and moving means mountedto said satellite for moving said at lest one surface member withrespect to said satellite whereby the movement of said at least onesurface member causes the center of gravity of said satellite to shiftfrom a first position to a second position such that said firstresultant force direction of the vector generated by said thruster meansacts through aid second position of said center of gravity.
 2. Theattitude control system according to claim 1 wherein said at least onesurface member is a plane solar generator panel extending along saidfirst direction, and wherein said attitude control system furthercomprises drive means mounted to said satellite and connected to saidplane solar generator panel for rotating said plane solar generatorpanel about said first direction.
 3. The attitude control systemaccording to claim 2 wherein said moving means comprises said drivemeans and further comprises a second drive means disposed between saiddrive means and said plane solar generator panel.
 4. The attitudecontrol system according to claim 3 wherein said second drive means is arotary motor having an axis inclined at an angle relative to said firstdirection.
 5. The attitude control system according to claim 4 whereinsaid angle is between about 2° and about 15°.
 6. The attitude controlsystem according to claim 1 wherein said moving means comprises a firstpivot motor having an axis which is transverse to said first direction.7. The attitude control system according to claim 6 wherein said firstpivot motor imparts to said at least one surface member a range ofmovement of at most about 15° relative to said first direction.
 8. Theattitude control system according to claim 6 wherein said moving meansfurther comprises a second pivot motor having an axis which istransverse to said first direction and which is inclined relative tosaid axis of said first pivot motor.
 9. The attitude control systemaccording to claim 1 wherein said moving means comprises:a deformablearticulated triangle coupled to said at least one surface member andsaid satellite; and a linear motor extending in a direction inclinedrelative to said first direction, said linear motor being mounted on oneside of said deformable articulated triangle.
 10. The attitude controlsystem according to claim 1 wherein said orbit control means compriseselectric thrusters.
 11. The attitude control system according to claim 1wherein said orbit control means comprises electric arc ionizationthrusters.
 12. The attitude control system according to claim 1 whereinsaid attitude stabilizing means during an orbit control maneuvercomprises said at least one surface member and said moving means. 13.The attitude control system according to claim 1 wherein said attitudestabilizing means during an orbit control maneuver comprises anorientable angular momentum system.
 14. A satellite adapted to bestabilized in attitude about roll, yaw and pitch axes in an at leastapproximately circular terrestrial orbit, said satellite comprising:abody having North and South sides; an attitude sensor system mounted tosaid body; an onboard computer in communication with said attitudesensor system; at least one solar generator panel member extendingsubstantially parallel to said pitch axis adapted to be exposed to solarradiation; means coupled to said body and said at least one solargenerator panel member for rotating said at least one solar generatorpanel member about said pitch axis, said rotating means being controlledby said onboard computer so as to remain at least approximatelyperpendicular to said solar radiation; a kinetic energy storage systemin communication with said onboard computer; an attitude control andorbit correction propulsion system in communication with said onboardcomputer, said attitude control and orbit correction propulsion systembeing exclusively electrical and comprising a first pair of electricthrusters disposed substantially symmetrically relative to a planeincluding said pitch and yaw axes, said first pair of electricalthrusters being inclined relative to a plane including said roll and yawaxes, an further being inclined relative to said plane including saidpitch and yaw axes and still further being inclined not more thanapproximately 20° relative to plane including said roll and pitch axessuch that a resultant force vector of said first pair of electricthrusters is in a first resultant force direction; and first meansmounted to said body for tilting said at least one solar generator panelmember relative to said pitch axis whereby the tilting of said at leastone solar generator panel member causes the center of gravity of saidsatellite to shift from a first position to a second position such thatsaid first resultant force direction of the vector generated by saidfirst pair of electric thrusters acts through said second position ofsaid center of gravity, said first tilting means being controlled bysaid onboard computer.
 15. A satellite according to claim 14 whereinsaid first pair of electric thrusters are disposed near one of saidNorth and South sides.
 16. A satellite according to claim 15 whereineach thruster of said first pair of electric thrusters is disposed nearan edge of said one of said North and South sides.
 17. A satelliteaccording to claim 16 wherein each thruster of said first pair ofelectric thrusters is disposed substantially at a mid portion of saidedge.
 18. A satellite according to claim 14 wherein said attitudecontrol and orbit correction propulsion system further comprises asecond pair of electric thrusters disposed substantially symmetricallyrelative to said plane including said pitch and yaw axes, said secondpair of electric thrusters being inclined relative to said planeincluding said roll and yaw axes in an opposite direction to said firstpair of electric thrusters, said second pair of electric thrustersfurther being inclined relative to said plane including said yaw andpitch axes and still further being inclined not more than approximately20° relative to said plane including said roll and pitch axes such thata resultant force vector of said first and second pair of electricthrusters is in a first resultant force direction.
 19. A satelliteaccording to claim 18 wherein said first and second pair of electricthrusters are substantially symmetrically disposed on opposite sides ofsaid plane including said roll and yaw axes.
 20. A satellite accordingto claim 18 wherein said attitude control and orbit correctionpropulsion system comprises only four electric thrusters.
 21. Asatellite according to claim 18 wherein each of said first and secondpairs of electric thrusters is inclined relative to said roll axis at anangle between about 40° and about 75° in absolute value and relative tosaid pitch axis at an angle between about 15° and about 65° in absolutevalue.
 22. A satellite according to claim 14 further comprising a secondsolar generator panel member mounted to said body so as to extendsubstantially parallel to said pitch axis in a direction substantiallyopposite to said at least one solar generator panel member, said secondsolar generator panel member being coupled to said body by a secondmeans for rotating said second solar generator panel member about saidpitch axis, said second solar generator panel member further beingcontrolled by said onboard computer so as to remain at leastapproximately perpendicular to said solar radiation, said second solargenerator panel being coupled to said body by a second means for tiltingsaid second solar generator panel member relative to said pitch axis,said first and second tilt means being controlled conjointly by saidonboard computer whereby the conjoint tilting of said at lest one andsaid second solar generator panel members causes the center of gravityof said satellite to shift from a first position to a second positionsuch that said first pair of electric thrusters acts through said secondposition of said center of gravity.
 23. A satellite according to claim14 wherein said kinetic energy storage system has no permanentgyroscopic stiffness, said kinetic energy storage system comprising atleast three reaction wheels whose angular momentum can re reduced tozero.
 24. A satellite according to claim 18 wherein at least one pair ofsaid first and second pairs of electric thrusters is a pair of ionthrusters.
 25. A satellite according to claim 18 wherein at least onepair of said first and second pairs of electric thrusters is a pair ofplasma thrusters.
 26. A satellite according to claim 14 furthercomprising:means mounted to said body for injection phase attitudecontrol as said satellite is injected into an operational orbit; and asingle-propellant storage tank connected to said attitude control meansand said injection phase first pair of electric thrusters.
 27. Asatellite according to claim 26 further comprising a dual propulsion andattitude control propellant system connected to said first pair ofelectric thrusters and said injection phase attitude control means. 28.A satellite according to claim 14 wherein said rotating means and saidfirst tilting means are controlled by said onboard computer so as togenerate attitude correction about said roll, yaw and pitch axes usingsolar radiation pressure on said at least one solar generator panelmember.
 29. A satellite according to claim 14 wherein said onboardcomputer controls yaw attitude correction of said satellite with saidfirst pair of electric thrusters.
 30. A method of controlling theattitude of a satellite in a roll, yaw and pitch direction, saidsatellite having at least one surface member exposed to solar radiationand extending from said satellite in a first direction away from saidsatellite, an attitude control device and a computer adapted tocommunicate with said attitude control device and said at least onesurface member to determine the value of a resultant force vector ofsaid attitude control device, said method comprising the stepsof:applying orbit control thrusts utilizing a pair of thrusters mountedto said satellite that the resultant thrust force vector is in a firstresultant force direction; and tilting said at least one surface memberwith respect to said pitch axis such that said at lest one surfacemember is titled to cause the center of gravity of said satellite toshift from a first position to a second position whereby said firstresultant force direction of said force vector generated by said pair ofthrusters acts through said second position of said center of gravity ofsaid satellite.
 31. A method according to claim 30 further comprisingthe steps of:conducting an orbit control maneuver wherein the attitudeof said satellite is stabilized by an orientable angular momentumsystem; and tilting said at least one surface member so as to besubstantially parallel to said solar radiation to stabilize saidsatellite in pitch and to move said orientable angular momentum systeminto a given orientation relative to said satellite.
 32. A methodaccording to claim 30 further comprising the step of applying orbitcontrol thrusts with electric thrusters which are aligned substantiallyparallel to a roll-pitch plane of said satellite.